1. Field of the Invention
The present invention concerns actuator systems for aerospace controls and other aerospace functions. In particular, the invention concerns an actuator system for operating the aerodynamic control surfaces of an aircraft such as the ailerons, the wing leading edge and trailing edge flaps and slats, i.e. the flaperons, the elevators, the spoilers and the rudder. The actuator system of the present invention may also be employed for controlling the airbrakes, arrester hook, flight re-fuelling probe, the undercarriage or pallets, and the doors and locks generally.
2. Discussion of Prior Art
Actuator systems for operating aircraft control surfaces and for performing other aerospace functions are well known. Early such systems were in the form of mechanical linkage and lever arrangements operated manually by means of a control column or rudder bar by which the pilot could directly adjust the deflection of the aerodynamic control surfaces, for example.
Then, as aircraft increased in size and speed, additional force was required to operate the aerodynamic control surfaces, and various developments ensued.
The first of these involved the mounting of devices known as control tabs on the aerodynamic control surfaces themselves. Such tabs were in the form of small auxiliary control surfaces operated by the pilot and arranged in use to generate automatically an aerodynamic force designed to assist in moving the main control surface in the manner desired. In this way, the control tabs served to amplify aerodynamically the pilot""s effort for application to the main control surface, or they served to reduce the resistance to movement of the main control surface, in situations or aircraft where direct operation by the pilot would otherwise be impractical.
There were, however, a number of drawbacks to the use of control tabs, the most significant being potential risks of control surface flutter at resonant frequencies and potential shock wave problems at aircraft velocities approaching sonic conditions. Flutter arises from a structural interaction between the aircraft frame and the control surface resulting in a relative oscillatory movement between them, which becomes exacerbated as resonant frequencies are approached.
Consequently, alternative actuator systems were developed for enabling the pilot to move the control surfaces, such systems constituting either power-assisted control systems or power-operated control systems. In both cases, the power was provided by hydraulic actuators, comprising pumps driven from the main aircraft engine(s) and hydraulic jacks or piston and cylinder arrangements connected by fluid lines to the relevant pumps. In a power assisted control system, the pilot still employs a control column or rudder bar for operating the control surface, and the hydraulic jacks are arranged in such a way as to assist his efforts. In a power-operated control system, the pilot simply operates switches and valves to actuate the hydraulic jacks which themselves actuate the control surfaces.
The hydraulic actuators employed in such power-assisted and power-operated control systems are capable of producing high forces, and the necessary hinge moments for deflecting the control surfaces can readily be generated within relatively small space requirements. However, hydraulic systems are prone to leakage and wear, and ensuring that they remain at all times in a satisfactory operating condition requires a high level of maintenance.
This has led to attempts to produce electrically powered actuation systems. A number of proposals have been put forward and are employed in limited applications, but hitherto no satisfactory electrically powered actuation system has yet been produced.
One currently available electrical arrangement features an electro-hydraulic actuator, comprising an electrical power source driven from the main aircraft propulsion engine(s) and a hydraulic actuator arranged to be operated by the electrical power source for deflecting the control surfaces. This system suffers from all of the disadvantages of the known hydraulic actuator systems, together with the additional problems that the power source has a high mass, is expensive to produce and requires a substantial amount of aircraft space.
Another electrical arrangement currently available features a conventional electric motor driven from the main aircraft propulsion engine(s) together with a gearbox/ball screw and mechanical linkage to provide the required force and stroke for operating the control surfaces. Such arrangements suffer from problems of wear and a tendency to jam, which is a significant drawback considering the safety requirements in aircraft flight.
A further difficulty in employing electrical actuation systems for aircraft control surfaces concerns the dissipation of heat, particularly for high load, high duty cycle devices, especially in the kinds of environment in which many actuators are required to function. By way of example, the flaperon actuators of a fast jet will typically be located in an unconditioned wing bay having an ambient temperature of the order of 90xc2x0 C. In hydraulic actuators, the oil employed for driving them can also be used for cooling purposes and the same pipework can serve for both functions. The majority of the heat generated is thus removed by the circulating hydraulic fluid and is simply cooled by an additional fuel cooled oil cooler. By contrast, an electric actuator requires a separate cooling system, including separate piping and cooling fluid as well as a cooler, and this adds to the bulk of the actuator system.
Current efforts to introduce electrical actuation systems for aircraft control surfaces have to a large extent been based simply on replacing the existing hydraulic actuators with electric actuators. However, as indicated above, this leads to a new set of problems. For a similar power and duty cycle, a change from hydraulic actuation to electrical actuation will result in:
Significantly increased actuator cooling problems
Increased space requirements
Increased actuator mass
Increased actuator cost.
Another problem concerns the power requirements for the currently proposed and available electrical actuators. Such an actuator when employed to operate the flaperons of a modern fast jet typically requires a power supply of the order of 270 volt DC and draws peak currents between 100 and 150 amperes. This poses a significant safety risk in passenger applications.
It is an aim of the present invention to provide an actuator system which overcomes the above problems.
Its is also an aim of the present invention to provide an actuator system employing an electrical actuator which is compact and reliable and in which cooling requirements and power consumption are reduced.
According to the present invention, there is provided an actuator system for use in an aircraft control or operating system, comprising:
control means operable in response to an input for generating a control signal;
an electrical actuator responsive to the control signal for operating an aircraft flight control surface or other aircraft apparatus; and
means for aerodynamically assisting the electrical actuator in order to reduce the load on the electrical actuator in use.
In a preferred form of the invention described below, the actuator system is employed for activating an aircraft flight control surface, and the electrical actuator constitutes a linear electric motor while the means for aerodynamically assisting the electrical actuator comprise a control tab incorporated in or mounted on the control surface.
As described, the linear electric motor may be mounted within the aircraft frame, that is within the fuselage or a wing, or it may be mounted in the control tab itself. The linear electric motor may be arranged to deflect the control tab directly or by means of a linkage mechanism, or it may be arranged to deflect servo tabs mounted on the control tab, again either directly or through a linkage mechanism.
Advantageously, the control means are arranged to receive feedback signals, for example relating to control surface position and acceleration, for generating an appropriate control signal to control tab acceleration. Such feedback signals may also represent the stroke position of the linear electric motor.
The present invention combines the advantages of a relatively low powered electrical actuator with aerodynamic amplification of, or an aerodynamic reduction in resistance to, the force required for operating the aircraft control surface or other apparatus. Potentially, this can enable very significant reductions in actuator mass and size, and in the cooling requirements and power consumption.
Linear electric motor technology offers simplicity, low wear, low maintenance and high accuracy. In particular, linear electric motors are capable of long strokes, fast response times, and a high degree of positional accuracy compared with existing aircraft actuators. Although the achievable output forces compared with known actuation techniques are relatively low, this is compensated by the use of aerodynamic amplification or other aerodynamic assistance, for example in the form of control tabs.